Layered fuselage shield

ABSTRACT

An aircraft defining a longitudinal direction and a lateral direction is provided. The aircraft includes: a fuselage; an engine mounted at a location spaced from the fuselage of the aircraft, the engine comprising a plurality of rotor blades; and a fuselage shield attached to or formed integrally with the fuselage at a location in alignment with the plurality of rotor blades along the lateral direction, the fuselage shield comprising a first layer defining a first density and a second layer defining a second density, the first density being different than the second density.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a non-provisional application claiming the benefitof priority under 35 U.S.C. § 119(e) to U.S. Provisional Application No.62/915,364, filed Oct. 15, 2019, which is hereby incorporated byreference in its entirety.

FIELD

This application is generally directed to armor for a fuselage of anaircraft and an aircraft including the same.

BACKGROUND

Traditionally, gas turbine engines used in aircraft, e.g., a turboprop,unducted or ducted fan, generally include a turbomachine and a rotorassembly, along with a nacelle surrounding the rotor assembly (which mayalso be referred to as a fan assembly). The turbomachine generallyincludes a high-pressure spool and a low speed spool. A combustionsection receives pressurized air, which is mixed with fuel and combustedwithin a combustion chamber to generate combustion gases. The combustiongases are provided first to a high-pressure turbine of the high-pressurespool, driving the high-pressure spool, and subsequently to a low speedturbine of the low speed spool, driving the low speed spool. The rotorassembly is typically coupled to the low speed spool and driven by thelow speed spool.

Blade loss may cause damage to the aircraft, especially where rotors ofthe engine are not enclosed within a nacelle or fan duct, which may atleast partially contain a thrown blade.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In an aspect of the present disclosure, an aircraft defining alongitudinal direction and a lateral direction is provided. The aircraftincludes: a fuselage; an engine mounted at a location spaced from thefuselage of the aircraft, the engine comprising a plurality of rotorblades; and a fuselage shield attached to or formed integrally with thefuselage at a location in alignment with the plurality of rotor bladesalong the lateral direction, the fuselage shield comprising a firstlayer defining a first density and a second layer defining a seconddensity, the first density being different than the second density.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic, cross-sectional view of a gas turbine engine inaccordance with an exemplary aspect of the present disclosure.

FIG. 2 is a schematic, cross-sectional view of a gas turbine engine inaccordance with another exemplary aspect of the present disclosure.

FIG. 3 is a schematic view of an aircraft incorporating a gas turbineengine in accordance with FIGS. 1 and/or 2 .

FIG. 5 is a schematic view of a fuselage shield in accordance with anexemplary embodiment of the present disclosure.

FIG. 5 is a cross-sectional view of a fuselage shield in accordance withanother exemplary embodiment of the present disclosure.

FIG. 6 is a cross-sectional view of a fuselage shield in accordance withyet another exemplary embodiment of the present disclosure.

FIG. 7 is a cross-sectional view of a fuselage shield in accordance withstill another exemplary embodiment of the present disclosure.

FIG. 8 is a perspective view of a portion of the exemplary fuselageshield of FIG. 7 .

FIG. 9 is a schematic view of a fuselage of the aircraft of FIG. 3including a first fuselage shield.

FIG. 10 is a schematic view of the fuselage of the aircraft of FIG. 3including a second fuselage shield.

FIG. 11 is a forward-looking-aft, cross-sectional view of the fuselageof the aircraft of FIG. 3 .

FIG. 12 is a forward-looking-aft, cross-sectional view of a fuselage ofan aircraft having a fuselage shield in accordance with an exemplaryembodiment of the present disclosure.

FIG. 13 is a forward-looking-aft, cross-sectional view of a fuselage ofan aircraft having a fuselage shield in accordance with anotherexemplary embodiment of the present disclosure.

FIG. 14 is a schematic, plan view of a fuselage of an aircraft includinga fuselage shield in accordance with the present disclosure.

FIG. 15 is a schematic view of the exemplary fuselage shield of FIG. 14.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention.

The word “exemplary” is used herein to mean “serving as an example,instance, or illustration.” Any implementation described herein as“exemplary” is not necessarily to be construed as preferred oradvantageous over other implementations.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle, and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like refer to bothdirect coupling, fixing, or attaching, as well as indirect coupling,fixing, or attaching through one or more intermediate components orfeatures, unless otherwise specified herein.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, and “substantially”, are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems.

For example, the approximating language may refer to being within a 1,2, 4, 10, 15, or 20 percent margin. These approximating margins mayapply to a single value, either or both endpoints defining numericalranges, and/or the margin for ranges between endpoints.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

The inventors of the present disclosure have found that certain benefitsmay be achieved by removing a nacelle from a traditional turbofanengine. The inventors of the present disclosure have also found that itmay be beneficial to include a fuselage shield to provide additionalprotection for the fuselage proximate the rotor blades of the rotorassembly. The inventors of the present disclosure have also found thatthere are multiple functions that may be needed to be performed by thefuselage shield, including, e.g., protection against debris piercing thefuselage, absorption of energy from the debris, and preventing theenergy from materially deforming the fuselage.

More specifically, the inventors of the present disclosure have foundthat a fuselage shield configured to effectively provide protectionagainst debris piercing the fuselage, absorption of energy from thedebris, and preventing the energy from materially deforming the fuselagewould be useful. In at least certain exemplary aspects of the presentdisclosure, such benefits may be provided for an aircraft having anunducted rotor engine mounted at a location spaced from a fuselage ofthe aircraft through incorporation of a layered fuselage shield. Thefuselage shield may be attached to or formed integrally with thefuselage at a location in alignment with a single stage of rotor bladesof the unducted rotor assembly along the lateral direction. The fuselageshield may include a first layer defining a first density and a secondlayer defining a second density, with the first density being differentthan the second density.

Additionally or alternatively, a thickness of the first layer may bedifferent than a thickness of the second layer.

In certain aspects, the first layer may be an energy distribution layer,and the second layer may be an energy absorption layer.

Additionally or alternatively, still, the fuselage shield may include aload spreading layer configured to be positioned adjacent to thefuselage of the aircraft.

Moreover, in certain exemplary embodiments, the second layer may behermetically sealed within an interior of the fuselage shield. Such mayenable the fuselage shield to be more easily added to an aircraft as astand-alone product.

Referring now to the Drawings, FIG. 1 shows an elevationalcross-sectional view of an exemplary embodiment of a gas turbine engine10 as may incorporate one or more inventive aspects of the presentdisclosure. In particular, the exemplary gas turbine engine 10 of FIG. 1is a configured as a turboprop engine defining an axial direction A, aradial direction R, and a circumferential direction C (see, e.g., FIG. 4, extending about the axial direction A). As is seen from FIG. 1 , theengine 10 includes an array of airfoils arranged around a centrallongitudinal axis 14 of engine 10, and more particularly includes anarray of rotor blades 16 arranged around the central longitudinal axis14 of engine 10. The rotor blades 16 are arranged in typically equallyspaced relation around the centerline 14, and each blade has a root 22and a tip 24 and a span defined therebetween. The rotor assembly 12further includes a hub 45 located forward of the plurality of rotorblades 16.

It will be appreciated, that as used herein, the term “rotor blades” isused generically to refer to any rotatable blade, typically of anairfoil shape, that is rotatable by the engine 10 to generate thrust orcompress air for the engine 10. For example, in the embodiment of FIG. 1, the rotor blades 16 are sometimes also referred to as propellerblades, whereas in the embodiment of FIG. 2 , discussed below, the rotorblades 16 are sometimes also referred to as fan blades.

Additionally, the engine 10 includes a turbomachine 30 having core (orhigh speed system) 32 and a low speed system. The core 32 generallyincludes a high-speed compressor 34, a high speed turbine 36, and a highspeed shaft 38 extending therebetween and connecting the high speedcompressor 34 and high speed turbine 36. The high speed compressor 34(or at least the rotating components thereof), the high speed turbine 36(or at least the rotating components thereof), and the high speed shaft38 may collectively be referred to as a high speed spool of the engine.Further, a combustion section 40 is located between the high speedcompressor 34 and high speed turbine 36. The combustion section 40 mayinclude one or more configurations for receiving a mixture of fuel andair, and providing a flow of combustion gasses through the high speedturbine 36 for driving the high speed spool.

The low speed system similarly includes a low speed turbine 42 (or powerturbine) and a low speed shaft 46 extending between and connecting lowspeed turbine 42 and the plurality of rotor blades 16. Morespecifically, as shown in the embodiment illustrated in FIG. 1 , the lowspeed turbine 42 rotates and transfers rotational energy to the rotorblades 16 through the low speed shaft 46.

Referring still to FIG. 1 , the turbomachine 30 is generally encased ina cowl 48. The cowl 48 defines at least in part an inlet 50 and anexhaust 52, and includes a turbomachinery flowpath 54 extending betweenthe inlet 50 and the exhaust 52.

It will be appreciated, however, that the exemplary gas turbine engineof FIG. 1 is provided by way of example only. In other exemplaryembodiments, the gas turbine engine may have any other suitableconfiguration, such as any other suitable unducted, or open rotor,configuration. For example, referring now to FIG. 2 , an elevationalcross-sectional view of a gas turbine engine 10 in accordance withanother exemplary embodiment of the present disclosure is provided. Inparticular, the exemplary gas turbine engine 10 of FIG. 2 is aconfigured as a single unducted rotor engine. The exemplary singleunducted rotor engine of FIG. 2 may be configured in a similar manner asthe exemplary turboprop engine of FIG. 1 . For example, the exemplarysingle unducted rotor engine of FIG. 2 generally defines an axialdirection A, a radial direction R, and a circumferential direction C(see, e.g., FIG. 3 , extending about the axial direction A). Moreover,the engine 10 includes an array of airfoils arranged around a centrallongitudinal axis 14 of engine 10, and more particularly includes anarray of rotor blades 16 arranged around the central longitudinal axis14 of engine 10.

Additionally, the engine 10 includes a turbomachine 30 having core (orhigh speed system) 32, a low speed system, and a combustion section 40located between a high speed compressor 34 and a high speed turbine 36of the core 32. Moreover, as with the embodiment above, the low speedsystem includes a low speed turbine 42 and a low speed shaft 46.However, for the embodiment shown, the low speed system further includesa low speed compressor or booster 44, with the slow speed shaft 46extending between and connecting the low speed compressor 44 and lowspeed turbine 42.

It will be appreciated that although the engine 10 is depicted with thelow speed compressor 44 positioned forward of the high speed compressor34, in certain embodiments the compressors 34, 44 may be in aninterdigitated arrangement. Additionally, or alternatively, although theengine 10 is depicted with the high speed turbine 36 positioned forwardof the low speed turbine 42, in certain embodiments the turbines 36, 42may similarly be in an interdigitated arrangement.

Also similar to the embodiment of FIG. 1 , for the embodiment of FIG. 2the turbomachine 30 is generally encased in a cowl 48 defining at leastin part an inlet 50 and an exhaust 52, and including a turbomachineryflowpath 54 extending between the inlet 50 and the exhaust 52. For theembodiment shown, the inlet 50 is an annular or axisymmetric 360 degreeinlet 50 providing a path for incoming atmospheric air to enter theturbomachinery flowpath 54 (and compressors 44, 34, combustion section40, and turbines 36, 42).

Moreover, the exemplary gas turbine engine 10 depicted in FIG. 2additionally includes a non-rotating vane assembly 18 positioned aft ofthe rotor assembly 12 (i.e., non-rotating with respect to the centralaxis 14), which includes an array of airfoils also disposed aroundcentral axis 14, and more particularly includes an array of vanes 20disposed around central axis 14. The rotor blades 16 are arranged intypically equally spaced relation around the centerline 14, and eachblade has a root 22 and a tip 24 and a span defined therebetween. Itwill be appreciated that the vanes 20 each have a root 26 and a tip 28and a span defined therebetween. The rotor assembly 12 further includesa hub 45 located forward of the plurality of rotor blades 16.

Briefly, the inlet 50 is located between the rotor blade assembly 12 andthe fixed or stationary vane assembly 18 for the embodiment shown. Sucha location may be advantageous for a variety of reasons, includingmanagement of icing performance as well as protecting the inlet 50 fromvarious objects and materials as may be encountered in operation. Itwill be appreciated, however, that in other embodiments the inlet 50 maybe positioned at any other suitable location, e.g., aft of the vaneassembly 18, arranged in a non-axisymmetric manner, etc.

Referring still to FIG. 2 , the vane assembly 18 extends from the cowl48 and is positioned aft of the rotor assembly 12. The vanes 20 of thevane assembly 18 may be mounted to a stationary frame or other mountingstructure and do not rotate relative to the central axis 14. Forreference purposes, FIG. 1 also depicts the forward direction with arrowF, which in turn defines the forward and aft portions of the system. Asshown in FIG. 1 , the rotor assembly 12 is located forward of theturbomachine 30 in a “puller” configuration, and the exhaust 52 islocated aft of the guide vanes 20. As will be appreciated, the vanes 20of the vane assembly 18 may be configured for straightening out anairflow (e.g., reducing a swirl in the airflow) from the rotor assembly12 to increase an efficiency of the engine 10. For example, the vanes 20may be sized, shaped, and configured to impart a counteracting swirl tothe airflow from the rotor blades 16 so that in a downstream directionaft of both rows of airfoils (e.g., blades 16, vanes 20) the airflow hasa greatly reduced degree of swirl, which may translate to an increasedlevel of induced efficiency. Further discussion regarding the vaneassembly 18 is provided below.

It will be appreciated, however, that the exemplary single rotorunducted engine depicted in FIG. 2 is by way of example only, and thatin other exemplary embodiments, the engine 10 may have any othersuitable configuration, including, for example, any other suitablenumber of shafts or spools, turbines, compressors, etc. Additionally, oralternatively, in other exemplary embodiments, any other suitable gasturbine engine may be provided. For example, in other exemplaryembodiments, the gas turbine engine may be a ducted turbofan engine, aturboshaft engine, a turboprop engine, turbojet engine, etc. Further,although the exemplary engine 10 depicted in FIG. 2 includes a singleunducted rotor, in other exemplary embodiments, the engine may includemultiple unducted rotors (e.g., a pair of counter-rotating rotors).

Referring now to FIG. 3 , a schematic view of an aircraft 100 includingan engine 10 in accordance with an exemplary aspect of the presentdisclosure is provided. The exemplary aircraft 100 depicted generallyincludes a fuselage 102, a first wing 104A extending outwardly from afirst side/port side of the fuselage 102, and a second wing 104Bextending outwardly from a second side/starboard side of the fuselage102. The exemplary aircraft 100 further includes a first engine 10Amounted to the first wing 104A in an under wing configuration and asecond engine 10B mounted to the second wing 104B in and under wingconfiguration. The first and second engines 10A, 10B may each beconfigured in a similar manner as the exemplary engine 10 of FIG. 1 .Accordingly, each of the first and second engines 10A, 10B include arotor assembly 12 rotatable about a centerline axis 14 of the respectiveengine.

As will be appreciated, the rotor assembly 12 of the first and secondengines 10A, 10B include a single stage of unducted rotor blades 16. Assuch, there is no nacelle or similar structures surrounding the stage ofrotor blades 16 to contain the rotor blades 16 in the event of afailure. Although not depicted, the first and second engines 10A, 10Bmay additionally include a stage of stationary guide vanes (similar tothe stage of guide vanes 18 of FIG. 1 ).

Accordingly, for the exemplary aircraft 100 depicted, the aircraft 100includes a first fuselage shield 106 attached to or formed integrallywith the first side/port side of the fuselage 102 proximate the rotorassembly 12 of the first engine 10A and a second fuselage shield 108attached to or formed integrally with the second side/starboard side ofthe fuselage 102 proximate to the rotor assembly 12 of the second engine10B.

More specifically, for the embodiment depicted, the fuselage 102includes a fuselage exterior 103, which is generally an exterior surfaceof the fuselage 102, defined by an outer layer of the fuselage 102(e.g., an outer layer of sheet metal, optionally including one or morecoatings). The first fuselage shield 106 is placed on the fuselageexterior 103 and secured thereto, and similarly, the second fuselageshield 108 is also placed on the fuselage exterior 103 and securedthereto. In such a manner, it will be appreciated that for theembodiment depicted the first and second fuselage shields 106, 108 arenot integrated into an original design of the fuselage 102 and/orlocated within an outer layer of the fuselage 102, and instead are addedas a supplement to the fuselage exterior 103 as described herein.

More specifically, referring particular to the first engine 10A, it willbe appreciated that the aircraft 100 generally defines a longitudinaldirection L1 and a lateral direction L2. The first fuselage shield 106is attached to the fuselage 102 at a location aligned with the pluralityof rotor blades 16 of the rotor assembly 12 of the first engine 10Aalong the longitudinal direction L1. Further, the first fuselage shield106 defines a length 110 along the longitudinal direction L1, and theplurality of rotor blades 16 of the rotor assembly 12 define a rotorassembly diameter 112. For the embodiment depicted, the length 110 ofthe first fuselage shield 106 is equal to at least about 50% of therotor assembly diameter 112, such as at least about 60% of the rotorassembly diameter 112, such as at least about 75% of the rotor assemblydiameter 112, such as up to about 800% of the rotor assembly diameter112.

Notably, the exemplary aircraft 100 depicted may generally be referredto as a “narrow body” aircraft having a single aisle extending along alength thereof. In certain embodiments, the fuselage defines a width 105along the lateral direction L2 of at least 80 inches, such as at least90 inches, such as at least 100 inches, such as at least 110 inches,such as at least 130 inches. However, in other embodiments, the aircraft100 may alternatively be configured as a “wide body” aircraft having amultiple aisles extending along a length thereof and wider fuselage 102,such as up to 400 inches wide or up to 350 inches wide or up to 300inches wide. In at least certain exemplary embodiments, the highestseating capacity of a narrow-body aircraft may be 295 passengers, whilea wide body aircraft may be able to accommodate between 250 and 600passengers. For example, two-abreast aircraft typically seat 4 to 19passengers, three-abreast typically seat 24 to 45 passengers,four-abreast aircraft typically seat 44 to 80 passengers, five-abreastaircraft typically seat 85 to 130 passengers, six-abreast aircrafttypically seat 120 to 230 passengers. By contrast, a regional aircraftis generally smaller than the narrow and wide body aircraft, capable ofshorter flight times and carrying less passengers and/or cargo. Forexample, typical regional aircraft are designed to fly 100 passengers orless.

For embodiments of the present disclosure, the aircraft may have acruise altitude engine and aircraft operation at or above Mach 0.5,approximately Mach 0.55 and approximately Mach 0.85, or between March0.75 and Mach 0.85 for a cruise altitude. For example, an aircraft mayhave a cruise speed at cruise altitude. In various embodiments, theengine is applied to a vehicle with a cruise altitude up toapproximately 65,000 ft. In certain embodiments, cruise altitude isbetween approximately 28,000 ft and approximately 45,000 ft. In stillcertain embodiments, cruise altitude is expressed in flight levels basedon a standard air pressure at sea level, in which a cruise flightcondition is between FL280 and FL650. In another embodiment, cruiseflight condition is between FL280 and FL450. In still certainembodiments, cruise altitude is defined based at least on a barometricpressure, in which cruise altitude is between approximately 4.85 psiaand approximately 0.82 psia based on a sea level pressure ofapproximately 14.70 psia and sea level temperature at approximately 59degree Fahrenheit. In another embodiment, cruise altitude is betweenapproximately 4.85 psia and approximately 2.14 psia. It should beappreciated that in certain embodiments, the ranges of cruise altitudedefined by pressure may be adjusted based on a different reference sealevel pressure and/or sea level temperature.

In such a manner, it will be appreciated that the aircraft 100 isgenerally configured to carry a higher number of passengers, crew,and/or cargo than smaller aircraft, such as typical regional turbopropaircraft. With these smaller aircraft, the rotor blades (e.g.,propellers) of the engines tend to be relatively small and light with arelatively small diameter, and the power delivered by the engine islower than the larger and more powerful engines used on narrow body andwide body class commercial passenger aircraft. Moreover, among thedisclosed embodiments of an unducted rotor engine the delivered powerand rotor or blade size may be significantly larger than for a turbopropsuch that a risk of catastrophic damage in a failure event may beconsidered higher. For example, the rotor or blade size may be from 11to 14 feet in diameter, or 11 feet, 12 feet, or between 12 and 14 feetin diameter. Taking into consideration these factors while, at the sametime, addressing the impact in terms of increased weight, cost, andcomplexity and decreased efficiency (e.g., due to aerodynamic drag)associated with a shielding, the inventors arrived at a design that isbelieved to strike the correct balance between safety and efficiency,replaceability and maintenance of the shielding. It will be appreciatedthat the embodiments of shielding disclosed herein address the uniquechallenges faced in respect to passenger and crew protection for narrowand wide body commercial passenger aircraft engine or aircraftdesigners.

By contrast with current narrow body and wide body aircraft, theinventors of the present disclosure have found that when desirable toincorporate an open rotor engine (as opposed to, e.g., a ducted turbofanengine), and particularly when desirable to incorporate an open rotorengine with rotor blades defining a relatively large diameter (such asat least six feet, such as at least eight feet, such as at least tenfeet, such as at least twelve feet, such as up to 22 feet; such anengine is depicted in the embodiment of FIG. 2 ) the benefit and need ofthe fuselage shielding described herein may weigh in favor of inclusionof the fuselage shielding despite the increased weight and minimized,but still present, additional drag on the aircraft 100. As will beappreciated from the description herein, the fuselage shieldingdisclosed may allow for an aircraft to incorporate an engine 10 withunducted rotor blades without having to redesign and/or reconstruct thefuselage 102 to incorporate the fuselage shielding the inventors havefound to be beneficial.

It will be appreciated, however, that although described above as beingapplied to narrow body and wide body aircraft, in other embodiments,aspects of the present disclosure may further be applied to regionalaircraft.

Further, referring now to FIG. 4 , it will be appreciated that the firstfuselage shield 106 defines a plurality of layers. More specifically,the first fuselage shield 106 includes a first layer and a second layer.In the embodiment shown, the first layer is an energy distribution layer116 and the second layer is an energy absorption layer 118. Theexemplary fuselage shield 106 depicted additionally includes a thirdlayer, which for the embodiment shown is a load spreading layer 114. Theload spreading layer 114 is positioned adjacent to the fuselage 102 andthe energy distribution layer 116 is spaced from the fuselage 102.Further, the energy absorption layer 118 is positioned between the loadspreading layer 114 and the energy absorption layer 118. Morespecifically, for the embodiment shown, the energy distribution layer116 is positioned closer to the stage of unducted rotor blades 16 of thefirst engine 10A than the energy absorption layer 118, and the energyabsorption layer 118 is positioned closer to the stage of unducted rotorblades 16 of the first engine 10A than the load spreading layer 114.

In the event of a failure of the rotor assembly 12 of the first engine10A, debris may strike the energy distribution layer 116, which mayprevent such debris from cutting through and penetrating the firstfuselage shield 106. The energy absorption layer 118 may absorb theenergy transferred from the energy distribution layer 116 from thedebris. Finally, the load spreading layer 114 may distribute energy fromthe energy absorption layer 118 across the fuselage 102 to prevent anydeformation of the fuselage 102.

In certain exemplary embodiments, the energy distribution layer 116 maybe a metal (such as one or more pieces of sheet metal), a Kevlar, acarbon fiber composite (e.g., such as a carbon fiber composite with apolymeric resin, such as epoxy), a ceramic (such as a ceramic plate orceramic fiber), a combination thereof, or other material capable ofstopping debris from penetrating therethrough. The load spreading layer114 may be a metal layer, a graphite or an epoxy layer, a carbon fibercomposite (e.g., such as a carbon fiber composite with a polymericresin, such as epoxy), etc.

Further, it will be appreciated that the energy absorption layer 118 maybe formed of any material capable of absorbing a desired amount ofenergy. For example, referring now to FIGS. 5 and 6 , close-up, crosssectional views are provided of two fuselage shields 106 in accordancewith two exemplary embodiments of the present disclosure. Each of thefuselage shields 106 may be configured in similar manner as theexemplary first fuselage shield 106 described above. Accordingly, thefuselage shields 106 of FIGS. 5 and 6 generally include a load spreadinglayer 114, an energy absorption layer 118, and an energy distributionlayer 116. The load spreading layer 114 and energy distribution layer116 may be formed of materials described above, or other suitablematerials. The energy absorption layer 118 in these embodiments areformed of a relatively low dense material, having a low soliditypercentage (a percent by volume that the material that is a solid (ascompared to air or other gas) under nominal stresses). For example, thematerial forming the energy absorption layers 118 in these embodimentsmay define a solidity percentage less than 75%, such as less than 60%,such as less than 50%, such as less than 40%, such as at least 10%, suchas at least 25%.

Referring particularly to FIG. 5 , it will be appreciated that theenergy distribution layer 116 is formed of a lattice structure. Thelattice structure includes a plurality of interwoven members 140, whichmay bend to allow the structure to compress and absorb energy in theevent the fuselage shield 106 is struck with debris. The latticestructure may be formed through a 3D printing process/additivemanufacturing process, or any other suitable formation process. Themembers 140 may be formed of a metal, a plastic, an elastomericmaterial, etc.

Referring particularly to FIG. 6 , it will be appreciated that theenergy distribution layer 116 is formed of a honeycomb structure. Thehoneycomb structure includes a plurality of members 142 connected toform a polygonal geometric pattern, and more specifically for theembodiment shown, to form a hexagonal geometric pattern. The members 142may be flexible to allow the material to compress and absorb energy inthe event the fuselage shield 106 is struck with debris. The honeycombstructure may be formed through a 3D printing process/additivemanufacturing process, or any other suitable formation process. Themembers 142 may be formed of a metal, a plastic, an elastomericmaterial, etc.

It will be appreciated, however, that in other exemplary embodiments,the energy distribution layer 116 may be formed of any other suitablematerial/structure. For example, in other embodiments, the energydistribution layer 116 may be formed of, e.g., a foam material, apolyurethane material, or any other suitable material capable ofabsorbing energy. Additionally, the energy distribution layer 116 may beformed of any other suitable structure capable of absorbing energy.

For example, referring now to FIGS. 7 and 8 , close-up views areprovided of a fuselage shield 106 in accordance with another exemplaryembodiment of the present disclosure. The fuselage shield 106 depictedmay also be configured in a similar manner as the exemplary firstfuselage shield 106 described above. Accordingly, the fuselage shield ofFIGS. 1 and 8 generally includes a load spreading layer 114, an energyabsorption layer 118, and an energy distribution layer 116. The loadspreading layer 114 and energy distribution layer 116 may be formed ofmaterials described above, or any other suitable material. However, theenergy absorption layer 118 of the embodiment shown is formed of aplurality of sheets 144 and plurality of spacers 146 positioned betweenadjacent sheets 144 of the plurality of sheets 144.

Specifically, referring first to FIG. 7 , providing a close-up,cross-sectional view of the energy distribution layer 116, it will beappreciated that the plurality of sheets 144 are spaced along athickness 150 of the fuselage shield 106 between the load spreadinglayer 114 and the energy distribution layer 116. Further, the pluralityof sheets 144 depicted includes at least two sheets 144, and morespecifically includes four sheets 144. The plurality of sheets 144 mayinclude up to twenty-five (25) sheets 144. The plurality of sheets 144may be formed of a sheet metal material, a metal alloy, a composite, orany other suitable material. For example, in certain embodiments, one ormore of the plurality of sheets 144 may be formed of a metal composite,a steel, a depleted uranium, a tungsten, a titanium, an Inconel, amolybdenum, an aluminum, a magnesium, an aluminum-lithium alloy,combinations thereof, etc.

Further, referring now particularly to FIG. 8 , providing a perspectiveview of section of a pair of adjacent sheets 144, it will be appreciatedthat between each pair of adjacent sheets 144, the energy absorptionlayer 118 includes a plurality of spacers 146. For example, the energyabsorption layer 118 may include at least two spacers 146, and up to onehundred (100) spacers 146, such as up to fifty (50) spacers 146, such asup to twenty-five (25) spacers 146. The spacers 146 may define a lengthto width ratio between 10:1 and 1:10, such as between 5:1 and 1:5, suchas between 2:1 and 1:2, such as equal to about 1:1. Moreover, referringalso to FIG. 7 , it will be appreciated that the spacers 146 aregenerally misaligned along the thickness 150 of the energy absorptionlayer 118. Specifically, for the embodiment shown, the spacers 146 in afirst gap 148A defined between adjacent sheets 144 do not overlap asviewed along a thickness 150 of the fuselage shield 106 with the spacers146 in a second gap 148B (positioned adjacent to the first gap 148A anddefined in part by a common sheet 144).

In such a manner, it will be appreciated that for the exemplary energyabsorption layer 118 depicted in FIGS. 7 and 8 , the sheets 144 maydeform to absorb energy transferred thereto from the energy distributionlayer 116.

Referring now back also to FIG. 4 , it will be appreciated from thediscussion hereinabove that the load spreading layer 114 defines a firstdensity, the energy absorption layer 118 defines a second density, andthe energy distribution layer 116 defines a third density. For theembodiment depicted, the first density and the third density are eachgreater than the second density of the energy absorption layer 118. Forexample, in certain exemplary embodiments, the first density, the thirddensity, or both may be at least about 20% greater than the seconddensity, such as at least about 50% greater, such as at least about 100%greater, such as at least about 200% greater, such as at least about500% greater, such as up to about 10,000% greater than the seconddensity.

Further, referring specifically to the callout Circle 4 in FIG. 4 , itwill be appreciated that the energy distribution layer 116 defines afirst thickness 150A, the energy absorption layer 118 defines a secondthickness 150B, and the load spreading layer 114 defines a thirdthickness 150C. As used herein, the term “thickness” refers to a maximumthickness of the component being described, such as a maximum thicknessof the layer of the fuselage shield 106.

For the embodiment show, the first thickness 150A is at least 0.05inches and up to 2.5 inches, such as at least 0.1 inches, such as atleast 0.5 inches, such as at least 0.75 inches, such as up to 2.25inches, such as up to 2 inches. Further for the embodiment show, thethird thickness 150C is at least 0.05 inches and up to 2.5 inches, suchas at least 0.1 inches, such as at least 0.5 inches, such as at least0.75 inches, such as up to 2.25 inches, such as up to 2 inches. Further,still, for the embodiment shown, the second thickness 150B is at least0.25 inches and up to 4 inches, such as at least 0.35 inches, such as atleast 0.5 inches, such as at least 1 inch, such as at least 2 inches,such as up to 3.75 inches, such as up to 3.5 inches.

In such a manner, it will be appreciated that the second thickness 150Bof the energy absorption layer 118 may be greater than the firstthickness 150A and third thickness 150C, such as at least about 50%greater, such as at least about 100% greater, such as up to about 10000%greater. Such may allow for the lower density of the energy absorptionlayer 118, which may allow for the energy absorption layer 118 to absorbmore energy in the event the fuselage shield 106 is struck with debris.

As will also be appreciated from the Figs. and the discussion herein, itmay be beneficial to configure and/or orient the fuselage shield for theestimated debris it is meant to protect the fuselage 102 from. Forexample, referring now to FIGS. 9 and 10 (and also back to FIG. 3 ), itwill be appreciated that in certain exemplary embodiments, the firstfuselage shield 106 may be configured in a different manner than thesecond fuselage shield 108. For example, it will be appreciated that, asindicated by the circular arrows in FIG. 3 (extending about centerlines14), the plurality of rotor blades 16 of the rotor assemblies 12 of thefirst and second engines 10A, 10B rotate in the same rotationaldirection, but given their location on opposing sides of the fuselage102, approach the fuselage 102 in a different manner. For example, therotor blades 16 of the rotor assembly 12 of the first engine 10Aapproach from below the aircraft 100, while the rotor blades 16 of therotor assembly 12 of the second engine 10B approach from above theaircraft 100. As such, an estimated strike from debris from the firstengine 10A may be different from an estimated strike from debris fromthe second engine 10B in the event of a failure of these engines.

Therefore, in at least certain exemplary embodiments, the first fuselageshield 106 is positioned asymmetrically relative to the second fuselageshield 108 relative to a reference plane 120 extending along thelongitudinal direction L1 and a vertical direction V of the aircraft100, through a center 152 of the fuselage 102 and of the aircraft 100.

Specifically, for the embodiment depicted, the first fuselage shield 106may be mounted to the fuselage 102 at a different position and/or in adifferent orientation than the second fuselage shield 108. Morespecifically, referring particularly to FIG. 9 , providing a plan viewof a first side/port side of the fuselage 102, and to FIG. 10 ,providing a plan view of a second side/starboard side of the fuselage102, it will be appreciated that the first fuselage shield 106 andsecond fuselage shield 108 may be mounted in a manner particular for thespecific engines, and in a manner to most appropriately accommodate anydebris that may result from such engines.

More specifically, still, as will be appreciated from FIGS. 9 and 10 ,the asymmetrical positioning of the first fuselage shield 106 and thesecond fuselage shield 108 relative to the reference plane 120 isderived from a vertical positioning of the first fuselage shield 106 andsecond fuselage shield 108. For example, the first fuselage shield 106defines a top 122 and a bottom 124 along the vertical direction V, andsimilarly, the second fuselage shield 108 defines a top 126 and a bottom128 along the vertical direction V. The top 122 of the first fuselageshield 106 is position higher along the vertical direction V than thetop 126 of the second fuselage shield 108, and as such is closer to atop portion 130 of the fuselage 102. Further, the bottom 124 of thesecond fuselage shield 108 is positioned lower along the verticaldirection V than the bottom 128 of the first fuselage shield 106, and assuch is closer to a bottom portion 132 of the fuselage 102.

Such a configuration may also be seen in FIG. 11 , providing aforward-looking-aft, cross-sectional view of the fuselage 102 of theaircraft of FIG. 3 . As described above, the fuselage 102 includes thefirst fuselage shield 106 and the second fuselage shield 108. The firstfuselage shield 106 includes the top 122 and the bottom 124, and thesecond fuselage shield 108 includes the top 126 and the bottom 128. Asis depicted in FIG. 11 , the first fuselage shield 106 assembly definesa first absolute positioning angle 154 relative to a top portion of avertical reference line extending through the center 152 of the fuselage102, or more specifically relative to a top portion of the referenceplane 120. Similarly, the second fuselage shield 108 assembly defines asecond absolute positioning angle 156 relative to the top of thevertical reference line extending through the center 152 of the fuselage102, or more specifically relative to the top portion of the referenceplane 120. As used herein, the term “absolute positioning angle” refersto the absolute value of an angle between two lines. Further, it will beappreciated that the absolute positioning angle of the first fuselageshield 106 is measured from a reference line 158 extending from thecenter 152 of the fuselage 102 and a center of the first fuselage shield106 (as measured along a circumference of the fuselage 102), andsimilarly the absolute positioning angle of the second fuselage shield108 is measured from a reference line 160 extending from the center 152of the fuselage 102 and a center of the second fuselage shield 108 (alsoas measured along a circumference of the fuselage 102).

For the embodiment shown, the difference between the first absolutepositioning angle 154 and the second absolute positioning angle 156 isat least five degrees and up to fifty (50_degrees. For example, incertain exemplary embodiments, the difference between the first absolutepositioning angle 154 and the second absolute positioning angle 156 maybe at least ten degrees, such as at least fifteen (15) degrees, such asup to forty-five (45) degrees, such as up to forty (40) degrees, such asup to thirty-five (35) degrees.

As will be appreciated for the embodiment depicted, the asymmetricpositioning of the first fuselage shield 106 and the second fuselageshield 108 may provide for a desired amount of protection for thefuselage 102, without requiring excess fuselage armor, which may lead toan overall heavier aircraft 100, and further may increase an airflowresistance of the aircraft 100.

Moreover, it will be appreciated that the first fuselage shield 106 andsecond fuselage shield 108 may be sized and/or arranged to provide thedesired coverage for the particular gas turbine engine in question. Forexample, referring still to FIG. 11 , will be appreciated that the firstfuselage shield 106 defines a coverage span is measured along acircumference of the fuselage 102. More specifically, it will beappreciated that the first fuselage shield 106 defines a coverage spanangle 162 between the top 122 and the bottom 124 of the first fuselageshield 106, as measured from the center 152 of the fuselage 102. Thecoverage span angle 162 for the first fuselage shield 106 is at leastforty-five (45) degrees and up to one hundred and eighty (180) degrees.For example, in certain exemplary embodiments, the coverage span angle162 for the first fuselage shield 106 may be at least about fifty-five(55) degrees and up to about one hundred and twenty (120) degrees. Itwill be appreciated that the second fuselage shield 108 may define asimilar coverage span angle.

In such manner, it will further be appreciated that the fuselage shield106 may define a surface area sufficient to provide the desired amountof coverage. In at least certain exemplary embodiments, the firstfuselage shield 106, the second fuselage shield 108, or both may definea surface area of at least 720 square inches and up to 15,000 squareinches. For example, the first fuselage shield 106, the second fuselageshield 108, or both may define a surface area of at least 1000 squareinches, such as at least 1200 square inches, such as up to 13,000 squareinches, such as up to 10,000 square inches.

As explained above, certain areas of the fuselage 102 may be moresusceptible to higher force impacts from debris than others due to,e.g., a proximity to the gas turbine engines, a rotational direction ofthe rotor blades 16 of the gas turbine engine, etc. In order to furtherprotect from such higher force impacts, without unnecessarily increasinga weight of the fuselage shield 106 and aircraft, in at least certainexemplary embodiments, the fuselage shield 106 may be designed toaccommodate different force impacts at various positions along thefuselage shield 106. For example, referring now to FIGS. 12 and 13 , incertain exemplary embodiments the fuselage shield 106 may include aplurality of zones arranged along a circumference of the fuselage 102such that an impact resistance of the fuselage shield 106 varies betweeneach of the adjacent zones along the circumference of the fuselage 102.

Referring particularly to FIG. 12 , a forward-looking-aft,cross-sectional view of a fuselage 102 including a fuselage shield 106in accordance with an exemplary embodiment of the present disclosure isprovided. The exemplary fuselage shield 106 depicted may be configuredin a similar manner as the exemplary first fuselage shield 106 describedabove, or may be configured in any other suitable manner.

The exemplary fuselage shield 106 of FIG. 12 includes a first zone 164having a first impact resistance and a second zone 166 having a secondimpact resistance, with the first zone 164 and the second zone 166arrange along the circumference of the fuselage 102. More specifically,for the embodiment of FIG. 12 , the fuselage shield 106 further includesa third zone 168 having a third impact resistance. The second zone 166and the third zone 168 are arranged on opposing sides of the first zone164 along the circumference of the fuselage 102. The first impactresistance of the first zone 164 of the fuselage shield 106 is greaterthan the second impact resistance of the second zone 166 and is alsogreater than the third impact resistance of the third zone 168.

Specifically, for the embodiment shown, the variance in impactresistance is due at least in part to a thickness of the plurality ofzones. More specifically, for the embodiment shown, the first zone 164defines a first thickness 170 that is greater than a second thickness172 of the second zone 166 and greater than a third thickness 174 of thethird zone 168. When the fuselage shield 106 is configured in accordancewith one or more the exemplary embodiments above, such as the embodimentof FIG. 4 wherein the fuselage shield 106 includes a plurality oflayers, the configuration of FIG. 12 may facilitate an outer layer(e.g., energy distribution layer 116), a middle layer (e.g., energyabsorption layer 118), and/or an inner layer (e.g., load spreading layer114) of the first zone 164 being thicker than the respective layer ofthe second zone 166 and/or third zone 168.

Notably, for the embodiment of FIG. 12 , the different zones of thefuselage shield 106 are formed integrally together as a single fuselageshield 106. However, in other embodiments, the fuselage shield 106 mayhave still other configurations. Further for the embodiment shown, itwill be appreciated that the first zone 164, the second zone 166 and thethird zone 168 each define respective span angles 176A, 176B, 176C(measured in the same way the span angle 164 of the first fuselageshield 106 is measured in FIG. 11 ). The span angles 176A, 176B, 176Care approximately equal to one another in FIG. 12 . In otherembodiments, however, the span angles 176A, 176B, 176C may vary betweenthe zones 164, 166, 168. For example, in certain exemplary embodiments,each of the zones 164, 166, 168 may define a span angle 176A, 176B, 176Cequal to at least 10% of the total span angle 162 of the first fuselageshield 106 and up to 80% of the total span angle 162 of the firstfuselage shield 106.

For example, referring now to FIG. 13 , a forward-looking-aft,cross-sectional view of a fuselage 102 including a fuselage shield 106in accordance with another exemplary embodiment of the presentdisclosure is provided. The exemplary fuselage shield 106 FIG. 13 againincludes a plurality of zones arrange along a circumference of thefuselage 102 with an impact resistance of the fuselage shield 106varying between the adjacent zones. In the embodiment shown, theplurality of zones includes at least four zones and up to 10 zones.Specifically, for the embodiment shown, the plurality of zones includesfive zones (a first zone 164, a second zone 166, a third zone 168, afourth zone 178, and a fifth zone 180). Although not labeled, athickness of the fuselage shield 106 varies between each of the adjacentzones (e.g., adjacent zones 164 and 166, and adjacent zones 166 and178). For the embodiment of FIG. 13 , each of the adjacent zones areconfigured as separate components which may be individually attached tothe fuselage 102, or alternatively, may be attached to one another priorto being attached the fuselage 102.

It will be appreciated that although the differing impact resistancesbetween adjacent zones if the embodiments of FIGS. 12 and 13 is providedat least in part by varying a thickness of the respective adjacentzones, in other embodiments, the differing impact resistances betweenadjacent zones may be provided in any other suitable manner. Forexample, in other embodiments the differing impact resistances betweenadjacent zones may be provided through material choices, layer relativethicknesses, a combination thereof, etc. In such a manner, it will beappreciated that in other exemplary embodiments, the fuselage shield 106may define a plurality of zones having differing impact resistancesbetween adjacent zones, while maintaining a relatively constantthickness between adjacent zones. Such may be desired for, e.g.,aerodynamic reasons, aesthetic reasons, etc.

As will be appreciated, inclusion of a fuselage shield 106 havingdifferent zones, with adjacent zones having different impactresistances, may allow for the fuselage shield 106 to be tailored to thecoverage desired/required for the specific gas turbine engine, includinga position of the gas turbine engine relative to through the fuselage102, a spacing of the gas turbine engine relative to the fuselage 102, arotational direction of a rotor assembly 12 of the gas turbine engine,etc. In such a manner, the fuselage shield 106 may not add more weight,air resistance, etc. than necessary to provide the desired/requiredimpact resistance to the fuselage 102.

Further, it will be appreciated that although the exemplary zonesdescribed above are arranged along a circumference of the fuselage 102,it will be appreciated that in certain exemplary embodiments, one ormore of the zones may also be arranged along the longitudinal directionL1 of the aircraft 100 (e.g., along the length 110 of the fuselageshield 106; see FIG. 3 ).

Moreover, it will be appreciated that although in certain exemplaryembodiments, the exemplary fuselage shields 106 described above may(except as otherwise described) be configured in a similar manner as theexemplary first fuselage shield 106 described above, in otherembodiments, the fuselage shield 106 may be configured in any othersuitable manner. For example, in certain exemplary embodiments, thefuselage shield 106 may not include three layers, and instead may onlyinclude two layers, or alternatively may include any other suitablenumber of layers (e.g., one, four, five, six or more).

Moreover, as noted above, the first fuselage shield 106 and secondfuselage shield 108 may be attached to, or formed integrally with, thefuselage 102. For example, in certain exemplary embodiments, the firstfuselage shield 106, the second fuselage shield 108, or both may bewelded to the fuselage 102 or otherwise irremovably formed integrallywith the fuselage 102.

However, in other exemplary embodiments, it will be appreciated that thefirst fuselage shield 106, the second fuselage shield 108, or both maybe removably coupled to the fuselage 102. For example, referring now toFIG. 14 , a plan view of a fuselage shield 106 in accordance with yetanother exemplary embodiment of the present disclosure is depictedattached to a fuselage 102. The fuselage shield 106 may be configured ina similar manner as the exemplary first fuselage shield 106 describedabove, or alternatively may be configured in any other suitable manner.

For the embodiment shown, the fuselage shield 106 is removably coupledto the fuselage 102. As with the embodiments above, the fuselage shield106 depicted in FIG. 14 may be removably coupled to the fuselage 102 ata location in alignment with a stage of rotor blades of an unductedrotor assembly of the gas turbine engine along a lateral direction ofthe aircraft.

For the embodiment shown, the fuselage shield 106 is removably coupledto the fuselage 102 using a plurality of mechanical fasteners 138.Additionally, the fuselage shield 106 includes a plurality of openings136 through which a respective plurality of mechanical fasteners 138 mayextend to couple the fuselage shield 106 to the fuselage 102.

More specifically, still, referring briefly to the callout Circle 14,the plurality of openings 136 of the fuselage shield 106 may beconfigured as a plurality of countersunk screw openings for receivingcorrespondingly shaped screws, or other mechanical fasteners, that arecountersunk into the fuselage shield 106 to reduce or eliminate anaerodynamic drag generated by the mechanical fasteners 138.

It will be appreciated, however, that in other exemplary embodiments,the fuselage shield 106 may be attached to the fuselage 102 in any othersuitable manner, using any other suitable mechanical fasteners 138 orother fastening means. For example, in other embodiments, the fuselageshield 106 may be attached entirely the plurality of mechanicalfasteners 138, such as one or more countersunk screws, bolts, etc., orthrough some combination of mechanical fasteners 138, features attachedto or formed with the fuselage 102, a glue, an epoxy, or the like.

Referring still to FIG. 14 , will be appreciated that the fuselageshield 106 defines a perimeter 182 (an area extending around anoutermost edge of the fuselage shield 106, closer to the outermost edgethan a center), and the fuselage shield 106 is removably coupled to thefuselage 102 with a plurality fasteners, or rather, the plurality ofmechanical fasteners 138, arranged in a density of at least one fastenerper inch and up to twenty-five (25) fasteners per inch. Such aconfiguration may ensure the fuselage shield 106 is coupled to thefuselage 102 in a manner that prevents airflow from passing between thefuselage shield 106 and the fuselage 102, creating excess drag on theaircraft.

Moreover, it will be appreciated that the fuselage shield 106 includesadditional features to reduce a drag on the aircraft when the fuselageshield 106 is coupled to the fuselage 102 of the aircraft. For example,referring now to FIG. 15 , a cross-sectional view of the exemplaryfuselage shield 106 and fuselage 102 of FIG. 14 is depicted. As shown,the fuselage shield 106 extends along the longitudinal direction L1between a forward end 184 and an aft end 186. The forward end 184 of thefuselage shield 106 defines a forward end taper 188 having a forward endtaper angle 190 with the fuselage 102 of at least one (1) degree and upto fifteen (15) degrees. Specifically, for the embodiment shown, theforward end taper angle 190 of the forward end 184 is less than or equalto seven (7) degrees. Similarly, for the embodiment shown, the aft end186 defines an aft end taper 192 having an aft end taper angle 194 withthe fuselage 102 of at least one (1) degree and up to fifteen (15)degrees, such as up to seven (7) degrees. Inclusion of the forward endtaper 188 and the aft end taper 192 may reduce an aerodynamic drag onthe aircraft by virtue of the inclusion of the fuselage shield 106.

Moreover, for exemplary embodiment depicted, the fuselage shield 106includes a plurality of layers. Specifically, for the embodiment shown,the fuselage shield 106 includes a first/outer layer (e.g., energydistribution layer 116), a second/middle layer (e.g., energy absorptionlayer 118), and a third/inner layer (e.g., load spreading layer 114).Further, as noted above, the exemplary fuselage shield 106 is removablycoupled to the fuselage 102 using a plurality of mechanical fasteners138. For the embodiment shown, the fuselage shield 106 is morespecifically removably coupled to the fuselage 102 through thefirst/outer layer using the mechanical fasteners 138. Notably, themechanical fasteners 138 further extend through the third/inner layer.

In such a manner, it will be appreciated that the outer layer, themiddle layer, and the inner layer may be coupled to one another usingmechanical fasteners 138. In addition, for the embodiment depicted, thefuselage shield 106 defines a joint 196 between the outer layer and theinner layer. For the embodiment shown, the joint 196 is a weld jointattaching the outer and inner layer to one another, and coupling thefuselage shield 196 together.

However, in other embodiments, the first layer, the middle layer, and athird layer may be coupled to another at least in part using any othersuitable means, such as through a mechanical clamp, a resin bondment, acompression wrap, a weld joint, a lamination, or a combination thereof.

Referring still to FIG. 15 , it will also be appreciated that for theembodiment shown, the middle layer is completely enclosed within aninterior 198 of the fuselage shield 106. More specifically, for theembodiment shown, the middle layer is completely enclosed between theouter layer and the inner layer, the interior 198 defined by the outerlayer and the inner layer. More specifically, still, for the embodimentshown, the middle layer is hermetically sealed within the interior 198of the fuselage shield 106, or rather, hermetically seal between theouter layer and the inner layer.

As will be appreciated, having the fuselage shield 106 configured suchthat the layers are coupled to one another, and further such that amiddle layer (e.g., an energy absorption layer 118) is hermeticallysealed within an interior of the fuselage shield 106, may furtherfacilitate the fuselage shield 106 being a removable fuselage shield106. More specifically, one or more of such features may enable thefuselage shield to be installed on an aircraft without requiringadditional process steps or integrations with the aircraft to ensure theenergy absorption layer 118 is hermetically sealed with respect to anexternal airflow. For example, such may be desirable when the energyabsorption layer 118 defines a relatively low solidity percentage, suchthat if not hermetically sealed, airflow may flow thereto causingadditional drag on the aircraft.

It will be appreciated, however, that in other exemplary embodiments,the fuselage shield 106 may be configured in any other suitable manner,and attached to the fuselage 102 in any other suitable manner. Forexample, the fuselage shield 106 may not be removably coupled to thefuselage 102 of the aircraft 100, and instead may be permanently coupledto the fuselage 102. For example, the fuselage shield 106 may be welded,epoxied, brazed, etc. to the fuselage 102.

Further aspects of the invention are provided by the subject matter ofthe following clauses:

An aircraft defining a longitudinal direction and a lateral direction,the aircraft comprising: a fuselage; a single unducted rotor enginemounted at a location spaced from the fuselage of the aircraft, thesingle unducted rotor engine comprising an unducted rotor assemblyhaving a single stage of rotor blades; and a fuselage shield attached toor formed integrally with the fuselage at a location in alignment withthe single stage of rotor blades of the unducted rotor assembly alongthe lateral direction.

The aircraft of one or more of these clauses, further comprising: a wingextending from the fuselage generally along the lateral direction,wherein the single unducted rotor engine is mounted to the wing.

The aircraft of one or more of these clauses, wherein the singleunducted rotor engine is a first single unducted rotor engine, whereinthe fuselage shield is a first fuselage shield attached to or formedintegrally with a first side of the fuselage, and wherein the aircraftfurther comprises: a second single unducted rotor engine comprising anunducted rotor assembly having a single stage of rotor blades; a firstwing extending from the first side of the fuselage generally along thelateral direction, wherein the first single unducted rotor engine ismounted to the first wing; a second wing extending from a second side ofthe fuselage generally along the lateral direction, wherein the secondsingle unducted rotor engine is mounted to the second wing; and a secondfuselage shield attached to or formed integrally with the second side ofthe fuselage at a second location in alignment with the single stage ofrotor blades of the second unducted rotor assembly along the lateraldirection.

The aircraft of one or more of these clauses, wherein the first fuselageshield is positioned asymmetrically to the second fuselage shieldrelative to a reference plane extending along the longitudinal directionand a vertical direction through a center of the aircraft.

The aircraft of one or more of these clauses, wherein the rotor assemblyof the first single unducted rotor engine and the rotor assembly of thesecond single unducted rotor engine each rotate in the same rotationaldirection, and wherein the first fuselage shield extends higher than thesecond fuselage shield along the vertical direction or lower than thesecond fuselage shield along the vertical direction.

The aircraft of one or more of these clauses, wherein the first fuselageshield defines a first absolute positioning angle relative to a top of avertical reference line extending through a center of the fuselage,wherein the second fuselage shield defines a second absolute positioningangle relative to the top of the vertical reference line, wherein adifference between the first absolute positioning angle and the secondabsolute positioning angle is greater than 5 degrees and less than fiftydegrees.

The aircraft of one or more of these clauses, wherein the fuselageshield defines a top end and a bottom end along a vertical direction ofthe aircraft, wherein the fuselage shield defines a coverage span anglebetween the top end and the bottom end as measured from a center of thefuselage of at least 45 degrees and up to about 180 degrees.

The aircraft of one or more of these clauses, wherein the fuselageshield comprises a first zone having a first impact resistance and asecond zone having a second impact resistance, wherein the first zoneand second zone are arranged along a circumference of the fuselage.

The aircraft of one or more of these clauses, wherein the fuselageshield further comprises a third zone having a third impact resistance,wherein the second zone and third zone are arranged on opposing sides ofthe first zone along the circumference of the fuselage, and wherein thefirst impact resistance is greater than the second impact resistance andgreater than the third impact resistance.

The aircraft of one or more of these clauses, wherein the fuselageshield comprises a plurality of zones arranged along a circumference ofthe fuselage, wherein an impact resistance of the fuselage shield variesbetween each of the adjacent zones.

The aircraft of one or more of these clauses, wherein a thickness of thefuselage shield varies between each of the adjacent zones.

The aircraft of one or more of these clauses, wherein the plurality ofzones includes at least 4 zones and up to 10 zones.

The aircraft of one or more of these clauses, wherein the fuselageshield defines a surface area between 720 square inches and 15,000square inches.

A fuselage shield assembly for use with a fuselage of an aircraft havinga single unducted rotor engine, the aircraft defining a longitudinaldirection and a lateral direction, the fuselage shield assemblycomprising: a body configured to be attached to or formed integrallywith the fuselage of the aircraft at a location in alignment with thesingle unducted rotor engine along the lateral direction, bodycomprising a plurality of zones configured to be arranged along acircumference of the fuselage when coupled to the fuselage of theaircraft, wherein an impact resistance of the fuselage shield variesbetween each of the adjacent zones.

The fuselage shield assembly of one or more of these clauses, whereinthe plurality of zones includes a first zone having a first impactresistance and a second zone having a second impact resistance, whereinthe first zone and second zone are configured to be arranged along acircumference of the fuselage.

The fuselage shield assembly of one or more of these clauses, whereinthe fuselage shield further comprises a third zone having a third impactresistance, wherein the second zone and third zone are arranged onopposing sides of the first zone, and wherein the first impactresistance is greater than the second impact resistance and greater thanthe third impact resistance.

The fuselage shield assembly of one or more of these clauses, wherein athickness of the fuselage shield varies between each of the adjacentzones.

The fuselage shield assembly of one or more of these clauses, whereinthe plurality of zones includes at least 4 zones and up to 10 zones.

The fuselage shield assembly of one or more of these clauses, whereinthe fuselage shield defines a surface area between 720 square inches and15,000 square inches.

The fuselage shield assembly of one or more of these clauses, whereinthe body of the fuselage shield assembly defines a top end and a bottomend along a vertical direction of the aircraft when coupled to thefuselage of the aircraft, wherein the body defines a coverage span anglebetween the top end and the bottom end as measured from a center of thefuselage of at least 45 degrees and up to about 180 degrees when coupledto the fuselage of the aircraft.

An aircraft defining a longitudinal direction and a lateral direction,the aircraft comprising: a fuselage; an unducted rotor engine mounted ata location spaced from the fuselage of the aircraft, the unducted rotorengine comprising an unducted rotor assembly having a stage of unductedrotor blades; and a fuselage shield removably coupled to the fuselage ata location in alignment with the stage of rotor blades of the unductedrotor assembly along the lateral direction.

The aircraft of one or more of these clauses, wherein the fuselageshield is removably coupled to the fuselage using a plurality ofmechanical fasteners.

The aircraft of one or more of these clauses, wherein the fuselageshield defines a perimeter, and wherein the fuselage shield is removablycoupled to the fuselage with a plurality of fasteners arranged in adensity of at least one fastener per inch and up to 25 fasteners perinch.

The aircraft of one or more of these clauses, wherein the fuselageshield defines a forward end and an aft end, wherein the forward enddefines a forward end taper angle of at least 1 degree and up to 15degrees.

The aircraft of one or more of these clauses, wherein the forward endtaper angle of the forward end is less than or equal to 7 degrees.

The aircraft of one or more of these clauses, wherein the aft enddefines an aft end taper angle of at least 1 degree and up to 15degrees.

The aircraft of one or more of these clauses, wherein the fuselageshield comprises a first layer and a second layer.

The aircraft of one or more of these clauses, wherein the second layeris hermetically sealed within an interior of the fuselage shield.

The aircraft of one or more of these clauses, wherein the fuselageshield is removably coupled to the fuselage through the first layer.

The aircraft of one or more of these clauses, wherein the first layer isan energy distribution layer, wherein the second layer is an energyabsorption layer.

The aircraft of one or more of these clauses, wherein the unducted rotorengine is a single unducted rotor engine, and wherein the unducted rotorassembly and the stage of unducted rotor blades are a single unductedrotor assembly and a single stage of unducted rotor blades,respectively.

The aircraft of one or more of these clauses, wherein the fuselageshield defines a surface area between 720 square inches and 15,000square inches.

The aircraft of one or more of these clauses, wherein the fuselageshield assembly defines a top end and a bottom end along a verticaldirection of the aircraft, wherein the fuselage shield defines acoverage span angle between the top end and the bottom end as measuredfrom a center of the fuselage of at least 45 degrees and up to about 180degrees.

An aircraft defining a longitudinal direction and a lateral direction,the aircraft comprising: a fuselage; an unducted rotor engine mounted ata location spaced from the fuselage of the aircraft, the unducted rotorengine comprising an unducted rotor assembly having a stage of unductedrotor blades; and a fuselage shield attached to or formed integrallywith the fuselage at a location in alignment with the single stage ofrotor blades of the unducted rotor assembly along the lateral direction,the fuselage shield comprising a first layer defining a first densityand a second layer defining a second density, the first density beingdifferent than the second density.

The aircraft of one or more of these clauses, wherein a thickness of thefirst layer is different than a thickness of the second layer.

The aircraft of one or more of these clauses, wherein the first layer isan energy distribution layer, and wherein the second layer is an energyabsorption layer.

The aircraft of one or more of these clauses, wherein the thickness ofthe first layer is at least 0.05 inches and up to 2.5 inches, andwherein the thickness of the second layer is at least 0.25 inches and upto 4 inches.

The aircraft of one or more of these clauses, further comprising a thirdlayer, wherein the third layer is a load spreading layer, and wherein athickness of the third layer is at least 0.05 inches and up to 2.5inches.

The aircraft of one or more of these clauses, wherein the first layer isformed of a Kevlar, a metal, a carbon fiber composite, a ceramic, or acombination thereof, and wherein the second layer comprises a honeycombstructure, a lattice structure, a foam material, a polyurethanematerial, or a combination thereof.

The aircraft of one or more of these clauses, wherein the first densityis greater than the second density.

The aircraft of one or more of these clauses, wherein the first densityis at least about 100% greater than the second density.

The aircraft of one or more of these clauses, wherein the energydistribution layer is positioned closer to the stage of unducted rotorblades than the energy absorption layer.

The aircraft of one or more of these clauses, wherein the fuselageshield further comprises a load spreading layer defining a thirddensity, wherein the third density is greater than the first density.

The aircraft of one or more of these clauses, wherein the energydistribution layer is positioned closer to the stage of unducted rotorblades than the energy absorption layer, and wherein the energyabsorption layer is positioned closer to the stage of unducted rotorblades than the load spreading layer.

The aircraft of one or more of these clauses, wherein the second layeris hermetically sealed within an interior of the fuselage shield.

The aircraft of one or more of these clauses, wherein the fuselageshield is coupled to the fuselage through the first layer.

The aircraft of one or more of these clauses, wherein the first layerand the second layer are secured to one another at least in part using amechanical clamp, a resin bondment, a compression wrap, a weld joint, alamination, or a combination thereof.

The aircraft of one or more of these clauses, wherein the first layer isan energy distribution layer, wherein the second layer is an energyabsorption layer, and wherein the second layer is formed of a pluralityof sheets and plurality of spacers positioned between adjacent sheets ofthe plurality of sheets.

A fuselage shield assembly for use with a fuselage of an aircraft havingan unducted rotor engine, the aircraft defining a longitudinal directionand a lateral direction, the fuselage shield assembly comprising: a bodyformed of a plurality of layers configured to be attached to or formedintegrally with the fuselage of the aircraft at a location in alignmentwith the unducted rotor engine along the lateral direction, theplurality of layers comprising a first layer and a second layer, thefirst layer defining a first density and the second layer defining asecond density, the first density being different than the seconddensity.

The fuselage shield assembly of one or more of these clauses, whereinthe first layer is an energy distribution layer, wherein the secondlayer is an energy absorption layer, and wherein the first density isgreater than the second density.

The fuselage shield assembly of one or more of these clauses, whereinthe first density is at least about 100% greater than the seconddensity.

The fuselage shield assembly of one or more of these clauses, whereinthe energy distribution layer is configured to be positioned closer tothe unducted rotor engine than the energy absorption layer.

The fuselage shield assembly of one or more of these clauses, whereinthe fuselage shield further comprises a load spreading layer defining athird density, wherein the third density is greater than the firstdensity, wherein the energy distribution layer is configured to bepositioned closer to the unducted rotor engine than the energyabsorption layer, and wherein the energy absorption layer is configuredto be positioned closer to the unducted rotor engine than the loadspreading layer.

An aircraft defining a longitudinal direction and a lateral direction,the aircraft comprising: a fuselage; an engine mounted at a locationspaced from the fuselage of the aircraft, the engine comprising rotorblades; and at least one fuselage shield removably coupled to thefuselage at a location in alignment with the rotor blades along thelateral direction.

The aircraft of one or more of these clauses, wherein the aircraft is anarrow body aircraft or a wide body aircraft.

The aircraft of one or more of these clauses, wherein the fuselage ofthe aircraft defines a width along a lateral direction of at least 80inches, such as at least 90 inches, such as at least 100 inches, such asat least 110 inches, such as at least 130 inches.

The aircraft of one or more of these clauses, wherein the fuselage ofthe aircraft defines a width along a lateral direction of up to 400inches, or up to 350 inches, or up to 300 inches.

The aircraft of one or more of these clauses, wherein the plurality ofrotor blades defines a diameter of at least six feet, such as at leasteight feet, such as at least ten feet, such as at least twelve feet.

The aircraft of one or more of these clauses, wherein the plurality ofrotor blades defines a diameter of up to 22 feet.

The aircraft of one or more of these clauses configured to carry morethan 100 passengers.

The aircraft of one or more of these clauses configured to carry morethan 150 passengers.

The aircraft of one or more of these clauses configured to carry lessthan 600 passengers.

The aircraft of one or more of these clauses having a cruise speedbetween Mach 0.5 and Mach 0.85.

The aircraft of one or more of these clauses having a cruise speedbetween Mach 0.75 and Mach 0.85.

The aircraft of one or more of these clauses having a cruise altitudebetween 28,000 feet and 65,000 feet.

The aircraft of one or more of these clauses having a cruise altitudebetween 28,000 feet and 45,000 feet.

The aircraft of one or more of these clauses having a cruise altitudeapproximately 4.85 psia and approximately 0.82 psia based on a sea levelpressure of approximately 14.70 psia and sea level temperature atapproximately 59 degree Fahrenheit.

The aircraft of one or more of these clauses having a cruise altitudeapproximately 4.85 psia and approximately 2.14 psia based on a sea levelpressure of approximately 14.70 psia and sea level temperature atapproximately 59 degree Fahrenheit.

An aircraft of one or more of these clauses incorporating a fuselageshield assembly of one or more of these clauses.

A fuselage shield assembly of one or more of these clauses incorporatedinto an aircraft of one or more of these clauses.

What is claimed is:
 1. An aircraft defining a longitudinal direction anda lateral direction, the aircraft comprising: a fuselage; an enginemounted at a location spaced from the fuselage of the aircraft, theengine comprising a plurality of rotor blades; and a fuselage shieldattached to or formed integrally with the fuselage at a location inalignment with the plurality of rotor blades along the lateraldirection, the fuselage shield comprising a first layer defining a firstdensity and a second layer defining a second density, the first densitybeing different than the second density.
 2. The aircraft of claim 1,wherein a thickness of the first layer is different than a thickness ofthe second layer.
 3. The aircraft of claim 1, wherein the first layer isan energy distribution layer, and wherein the second layer is an energyabsorption layer.
 4. The aircraft of claim 3, wherein the thickness ofthe first layer is at least 0.05 inches and up to 2.5 inches, andwherein the thickness of the second layer is at least 0.25 inches and upto 4 inches.
 5. The aircraft of claim 4, further comprising a thirdlayer, wherein the third layer is a load spreading layer, and wherein athickness of the third layer is at least 0.05 inches and up to 2.5inches.
 6. The aircraft of claim 3, wherein the first layer is formed ofa Kevlar, a metal, a carbon fiber composite, a ceramic, or a combinationthereof, and wherein the second layer comprises a honeycomb structure, alattice structure, a foam material, a polyurethane material, or acombination thereof.
 7. The aircraft of claim 3, wherein the firstdensity is greater than the second density.
 8. The aircraft of claim 3,wherein the first density is at least about 100% greater than the seconddensity.
 9. The aircraft of claim 3, wherein the energy distributionlayer is positioned closer to the plurality of rotor blades than theenergy absorption layer.
 10. The aircraft of claim 3, wherein thefuselage shield further comprises a load spreading layer defining athird density, wherein the third density is greater than the firstdensity.
 11. The aircraft of claim 10, wherein the energy distributionlayer is positioned closer to the plurality of rotor blades than theenergy absorption layer, and wherein the energy absorption layer ispositioned closer to the plurality of rotor blades than the loadspreading layer.
 12. The aircraft of claim 1, wherein the second layeris hermetically sealed within an interior of the fuselage shield. 13.The aircraft of claim 12, wherein the fuselage shield is coupled to thefuselage through the first layer.
 14. The aircraft of claim 1, whereinthe first layer and the second layer are secured to one another at leastin part using a mechanical clamp, a resin bondment, a compression wrap,a weld joint, a lamination, or a combination thereof.
 15. The aircraftof claim 1, wherein the first layer is an energy distribution layer,wherein the second layer is an energy absorption layer, and wherein thesecond layer is formed of a plurality of sheets and plurality of spacerspositioned between adjacent sheets of the plurality of sheets.
 16. Theaircraft of claim 1, wherein the aircraft is a narrow body or wide bodycommercial passenger aircraft.
 17. A fuselage shield assembly for usewith a fuselage of an aircraft having an unducted rotor engine, theaircraft defining a longitudinal direction and a lateral direction, thefuselage shield assembly comprising: a body formed of a plurality oflayers configured to be attached to or formed integrally with thefuselage of the aircraft at a location in alignment with the unductedrotor engine along the lateral direction, the plurality of layerscomprising a first layer and a second layer, the first layer defining afirst density and the second layer defining a second density, the firstdensity being different than the second density.
 18. The fuselage shieldassembly of claim 17, wherein the first layer is an energy distributionlayer, wherein the second layer is an energy absorption layer, andwherein the first density is greater than the second density.
 19. Thefuselage shield assembly of claim 18, wherein the first density is atleast about 100% greater than the second density.
 20. The fuselageshield assembly of claim 18, wherein the energy distribution layer isconfigured to be positioned closer to the unducted rotor engine than theenergy absorption layer.
 21. The fuselage shield assembly of claim 18,wherein the fuselage shield further comprises a load spreading layerdefining a third density, wherein the third density is greater than thefirst density, wherein the energy distribution layer is configured to bepositioned closer to the unducted rotor engine than the energyabsorption layer, and wherein the energy absorption layer is configuredto be positioned closer to the unducted rotor engine than the loadspreading layer.